Tip clearance measurement of a rotary wing aircraft

ABSTRACT

An aircraft is provided including an airframe, an extending tail, and a counter rotating, coaxial main rotor assembly including an upper rotor assembly with an upper blade and a lower rotor assembly with a lower blade. A first antenna in one of upper blade and the lower blade, and a second antenna in the other of the upper blade and the lower blade. An oscillator to apply an excitation signal to the first antenna. A blade proximity monitor to monitor a magnitude of the excitation signal and an output signal from the second antenna to determine a distance between the upper blade and the lower blade.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application claims the benefit of U.S. provisional patentapplication Ser. No. 62/058,424, filed Oct. 1, 2014, the entire contentsof which are incorporated herein by reference.

BACKGROUND

The subject matter disclosed herein relates generally to rotary wingaircraft and, more particularly, to a dual rotor, rotary wing aircraft.

Aircraft typically include an airframe and a number of aerodynamicsurfaces that generate lift. Rotary wing aircraft include a main rotorassembly comprising a number of rotor blades that generate lift andthrust for the aircraft. Coaxial rotary wing aircraft typically includetwo rotor assemblies positioned one on top of the other with space inbetween. Monitoring blade tip clearance on a coaxial rotary wingaircraft is important to prevent collision of the upper rotor assemblyblades and the lower rotor assembly blades.

BRIEF DESCRIPTION

In one exemplary embodiment, an aircraft includes an airframe; anextending tail; a counter rotating, coaxial main rotor assemblyincluding an upper rotor assembly with an upper blade and a lower rotorassembly with a lower blade; a first antenna in one of upper blade andthe lower blade, and a second antenna in the other of the upper bladeand the lower blade; an oscillator to apply an excitation signal to thefirst antenna; and a blade proximity monitor to monitor a magnitude ofthe excitation signal and an output signal from the second antenna todetermine a distance between the upper blade and the lower blade.

In addition to one or more of the features described above or below, oras an alternative, further embodiments could include wherein the bladeproximity monitor is mounted in a rotating portion of the main rotorassembly.

In addition to one or more of the features described above or below, oras an alternative, further embodiments could include wherein the bladeproximity monitor is configured to determine the azimuth location of theblades where the distance between the upper blade and the lower blade isdetermined.

In addition to one or more of the features described above or below, oras an alternative, further embodiments could include wherein the bladeproximity monitor is configured to transmit at least one of the distancebetween the upper blade and the lower blade and the azimuth location toat least one of an instrument system and a pilot display.

In addition to one or more of the features described above or below, oras an alternative, further embodiments could include wherein the bladeproximity monitor is configured to receive an RPM signal of the mainrotor assembly from a contactor.

In addition to one or more of the features described above or below, oras an alternative, further embodiments could include wherein theantennas are electric field antennas.

In addition to one or more of the features described above or below, oras an alternative, further embodiments could include wherein theantennas are magnetic field antennas.

In addition to one or more of the features described above or below, oras an alternative, further embodiments could include wherein theantennas are composed of a ferrite core material in the form of coilsaround the blade spar or embedded in the plane of the blade skin.

In addition to one or more of the features described above or below, oras an alternative, further embodiments could include wherein theantennas are composed of a small diameter magnet wire in the form ofcoils around the blade spar or embedded in the plane of the blade skin.

In addition to one or more of the features described above or below, oras an alternative, further embodiments could include wherein the bladesare rigidly connected to the upper rotor assembly and the lower rotorassembly.

In another exemplary embodiment, a method of operating an aircraft isprovided including generating, using an oscillator, an excitationsignal; transmitting, the excitation signal through a first antenna inone of an upper blade and a lower blade; receiving, the excitationsignal through a second antenna in the other of the upper blade and thelower blade and transmitting an output signal; and determining, with ablade proximity monitor, the distance between the upper blade and lowerblade by using the magnitude of the excitation signal and an outputsignal from the second antenna.

In addition to one or more of the features described above or below, oras an alternative, further embodiments could include wherein the bladeproximity monitor is mounted in a rotating portion of the main rotorassembly.

In addition to one or more of the features described above or below, oras an alternative, further embodiments could include wherein the bladeproximity monitor determines the azimuth location of the blades wherethe distance between the upper blade and the lower blade is determined.

In addition to one or more of the features described above or below, oras an alternative, further embodiments could include wherein the bladeproximity monitor transmits at least one of the distance between theupper blade and the lower blade and the azimuth location to at least oneof an instrument system and a pilot display.

In addition to one or more of the features described above or below, oras an alternative, further embodiments could include wherein the bladeproximity monitor receives an RPM signal of the main rotor assembly froma contactor.

In addition to one or more of the features described above or below, oras an alternative, further embodiments could include wherein theantennas are electric field antennas.

In addition to one or more of the features described above or below, oras an alternative, further embodiments could include wherein theantennas are magnetic field antennas.

In addition to one or more of the features described above or below, oras an alternative, further embodiments could include wherein theantennas are composed of a ferrite core material in the form of coilsaround the blade spar or embedded in the plane of the blade skin.

In addition to one or more of the features described above or below, oras an alternative, further embodiments could include wherein theantennas are composed of a small diameter magnet wire in the form ofcoils around the blade spar or embedded in the plane of the blade skin.

In addition to one or more of the features described above or below, oras an alternative, further embodiments could include wherein the bladesare rigidly connected to the upper rotor assembly and the lower rotorassembly.

BRIEF DESCRIPTION OF THE DRAWINGS

Referring now to the drawings wherein like elements are numbered alikein the several FIGURES:

FIG. 1 depicts a rotary wing aircraft in an exemplary embodiment;

FIG. 2 is a perspective view of a rotary wing aircraft in an exemplaryembodiment;

FIG. 2A depicts a planform of a rotor blade in an exemplary embodiment;

FIG. 3 is a perspective view of a gear train for a rotary wing aircraftin an exemplary embodiment;

FIGS. 3A and 3B depict power distribution in the gear box in hover andcruise modes in exemplary embodiments;

FIG. 4 is a perspective view of a gearbox and translational thrustsystem in an exemplary embodiment;

FIG. 5 is a perspective view of a rotor hub fairing in an exemplaryembodiment;

FIG. 6 depicts a flight control system in an exemplary embodiment;

FIG. 6A depicts a blade proximity detection system in an exemplaryembodiment;

FIG. 7 depicts a flight maneuver in an exemplary embodiment;

FIG. 8 depicts front, side and top views of an aircraft in an exemplaryembodiment;

FIG. 9 depicts an active vibration control (AVC) system in an exemplaryembodiment;

FIGS. 10 and 11 illustrate force vectors in exemplary hover states;

DETAILED DESCRIPTION

FIG. 1 depicts an exemplary embodiment of a rotary wing, verticaltakeoff and land (VTOL) aircraft 10. The aircraft 10 includes anairframe 12 with an extending tail 14. A dual, counter rotating, coaxialmain rotor assembly 18 is located at the airframe 12 and rotates about amain rotor axis, A. In an exemplary embodiment, the airframe 12 includestwo seats for flight crew (e.g., pilot and co-pilot) and six seats forpassengers. However an airframe 12 having another configuration iswithin the scope of the present disclosure. The main rotor assembly 18is driven by a power source, for example, one or more engines 24 via agearbox 26. The main rotor assembly 18 includes an upper rotor assembly28 driven in a first direction (e.g., counter-clockwise) about the mainrotor axis, A, and a lower rotor assembly 32 driven in a seconddirection (e.g., clockwise) about the main rotor axis, A, opposite tothe first direction (i.e., counter rotating rotors). Each of the upperrotor assembly 28 and the lower rotor assembly 32 includes a pluralityof rotor blades 36 secured to a rotor hub 38. In some embodiments, theaircraft 10 further includes a translational thrust system 40 located atthe extending tail 14 to provide translational thrust (forward orrearward) for aircraft 10.

Any number of blades 36 may be used with the rotor assembly 18. FIG. 2Adepicts a planform of a rotor blade 36 in an exemplary embodiment. Therotor assembly 18 includes a rotor hub fairing 37 generally locatedbetween and around the upper and lower rotor assemblies such that therotor hubs 38 are at least partially contained therein. The rotor hubfairing 37 provides drag reduction. Rotor blades 36 are connected to theupper and lower rotor hubs 38 in a hingeless manner, also referred to asa rigid rotor system. Although a particular aircraft configuration isillustrated in this non-limiting embodiment, other rotary-wing aircraftwill also benefit from embodiments of the invention. Although, the dualrotor system is depicted as coaxial, embodiments include dual rotoraircraft having non-coaxial rotors.

The translational thrust system 40 includes a propeller 42 connected toand driven by the engine 24 via the gearbox 26. The translational thrustsystem 40 may be mounted to the rear of the airframe 12 with atranslational thrust axis, T, oriented substantially horizontal andparallel to the aircraft longitudinal axis, L, to provide thrust forhigh-speed flight. The translational thrust axis, T, corresponds to theaxis of rotation of propeller 42. While shown in the context of apusher-prop configuration, it is understood that the propeller 42 couldalso be more conventional puller prop or could be variably facing so asto provide yaw control in addition to or instead of translationalthrust. It should be understood that any such system or othertranslational thrust systems may alternatively or additionally beutilized. Alternative translational thrust systems may include differentpropulsion forms, such as a jet engine.

Referring to FIG. 2, translational thrust system 40 includes a propeller42 and is positioned at a tail section 41 of the aircraft 10. Propeller42 includes a plurality of blades 47. In exemplary embodiments, thepitch of propeller blades 47 may be altered to change the direction ofthrust (e.g., forward or rearward). The tail section 41 includes activeelevators 43 and active rudders 45 as controllable surfaces.

Shown in FIG. 3 is a perspective view of portions of main rotor assembly18 and gearbox 26. The gearbox 26 includes an upper bull gear 44, whichrotates about the main rotor axis, A, and connected to the lower rotorassembly 32 via a lower rotor shaft 46 extending along the main rotoraxis, A. A lower bull gear 48, which rotates about the main rotor axis,A, and is connected to the upper rotor assembly 28 via an upper rotorshaft 50 extending along the main rotor axis, A, and through an interiorof the lower rotor shaft 46. Torque and rotational speed are provided tothe gearbox 26 via input shaft 52 that transmits the torque androtational speed from the engine(s) 24 to an input bevel 54 disposed atan input bevel shaft 56 of the gearbox 26 via an input bevel pinion 104.In some embodiments, the input bevel shaft 56 rotates about an inputbevel shaft axis 58 parallel to the main rotor axis A. The propeller 42is driven by a propeller output shaft 106 driven by a propeller outputgear 62 disposed at a quill shaft 102, or an extension of input bevelshaft 56. Transfer from the propeller output gear 62 is achieved viaconnection with a propeller output pinion 60 at the propeller outputshaft 106. To transfer torque from the input bevel shaft 56 to the lowerrotor assembly 32 and the upper rotor assembly 30, the gearbox 26includes a torque split gear reduction stage 64. The torque split gearreduction stage 64 splits torque from the input shaft 52 and applies thedivided torque to bull gears 44 and 48, respectively. While shown withthe propeller output shaft 106 driven by the propeller output gear 62,it is understood that such elements could be removed where the propeller42 is not used or is separately driven.

FIG. 3A illustrates power distribution through gearbox 26 to main rotorassembly 18 and propeller output shaft 106 during hover mode. In hover,power flows to torque split section to drive main rotor assembly 18. Thepropeller output shaft 106 spins at all times to drive features onpropeller box while propeller 42 is unclutched. During hover mode, themajority of power flows to the main rotor assembly 18.

FIG. 3B illustrates power distribution through gearbox 26 to main rotorassembly 18 and propeller output shaft 106 during cruise mode. In highspeed cruise, the majority of power flows to the propeller output shaft106 while the main rotor assembly 18 is operating near an autorotativestate.

Referring to FIG. 4, the main rotor assembly 18 is driven about the axisof rotation, A, through a main gearbox (MGB) 26 by a multi-enginepowerplant system 24, having two engine packages ENG1, ENG2 in theexample in FIG. 4. Although FIG. 4 depicts two engines 24, it isunderstood that aircraft 10 may use a single engine 24, or any number ofengines. The multi-engine powerplant system 24 generates power availablefor flight operations and couples such power to the main rotor assembly18 and the translational thrust system 40 through the MGB 26. The MGB 26may be interposed between the powerplant system 24, the main rotorassembly 18 and the translational thrust system 40.

A portion of the drive system, such as downstream of the MGB 26 forexample, includes a gearbox 90 (also referred to as a clutch). Thecombined gearbox 90 selectively operates as a clutch and a brake foroperation of the translational thrust system 40 with the MGB 26. Thegearbox 90 also operates to provide a rotor brake function for the mainrotor assembly 18.

The combined gearbox 90 generally includes an input 92 and an output 94generally defined along an axis parallel to rotational axis, T. Theinput 92 is generally upstream of the combined gearbox 90 relative theMGB 26 and the output 94 is downstream of the combined gearbox 90 andupstream of the pusher propeller system 40 (FIG. 2). The combinedgearbox 90 may be categorized by the technique used to disengage-engage(e.g., clutch) or stop (e.g., brake) the load such as friction,electromagnetic, mechanical lockup, etc., and by the method used toactuate such as mechanical, electric, pneumatic, hydraulic,self-activating, etc. It should be understood that various combinedgearbox 90 systems may be utilized to include but not to be limited tomechanical, electrically, hydraulic and various combinations thereof.

Referring to FIG. 5, an exemplary rotor hub fairing 37 is shown. Rotorhub fairing 37 is illustrated having generally elliptical, incross-section, upper and lower hub fairings 111 and 112, and anairfoil-type shape (in horizontal cross-section) for the shaft fairing103. The airfoil shape of the shaft fairing 103 includes a leading edge114, and a trailing edge 115 aft of the upper and lower fairings 111,112. The airfoil shape of the shaft fairing 103 additionally includes achord (not shown) that connects the leading and trailing edges 114, 115of the airfoil. In one embodiment, the airfoil shape, including theupper surface 116 and the lower surface 117, is symmetrical about aplane extending along the length of the shaft fairing 103 and containingthe axis of rotation, A. As noted above, the upper and lower rotor hubs38 may be positioned, at least partially, in the upper and lowerfairings 111, 112.

Portions of the aircraft 10 are controlled by a flight control system120 illustrated in FIG. 6. In one embodiment, the flight control system120 is a fly-by-wire (FBW) control system. In a FBW control system thereis no direct mechanical coupling between a pilot's controls and movablecomponents of aircraft 10. Instead of using mechanical linkages, a FBWcontrol system includes a plurality of sensors 122 which can sense theposition of controlled elements and generate electrical signalsproportional to the sensed position. The sensors 122 may also be useddirectly and indirectly to provide a variety of aircraft state data to aflight control computer (FCC) 124. The FCC 124 may also receive inputs126 as control commands from various sources. For instance, the inputs126 can be pilot inputs, auto-pilot inputs, navigation system basedinputs, or any control inputs from one or more control loops executed bythe FCC 124 or other subsystems. In response to inputs from the sensors122 and inputs 126, the FCC 124 transmits signals to various subsystemsof the aircraft 10.

Flight control system 120 may include a rotor interface 128 configuredto receive commands from the FCC 124 and control one or more actuators,such as a mechanical-hydraulic or electric actuators, for the upperrotor assembly 28 and lower rotor assembly 32. In an embodiment, inputs126 including cyclic, collective, pitch rate, and throttle commands thatmay result in the rotor interface 128 driving the one or more actuatorsto adjust upper and lower swashplate assemblies (not depicted) for pitchcontrol of the upper rotor assembly 28 and lower rotor assembly 32.Alternatively, pitch control can be performed without a swashplateassemblies using individual blade control (IBC) in the upper rotorassembly 28 and lower rotor assembly 32. The rotor interface 128 canmanipulate the upper rotor assembly 28 and lower rotor assembly 32independently. This allows different collective and cyclic commands tobe provided to the upper rotor assembly 28 and lower rotor assembly 32.

Flight control system 120 may include a translational thrust interface130 configured to receive commands from the FCC 124 to control one ormore actuators, such as a mechanical-hydraulic or electric actuators,for the control of the translational thrust system 40. In an embodiment,inputs 126 may result in the translational thrust interface 130controlling speed of propeller 42, altering the pitch of propellerblades 47 (e.g., forward or rearward thrust), altering the direction ofrotation of propeller 42, controlling gearbox 90 to employ a clutch 91to engage or disengage the propeller 42, etc.

Flight control system 120 may include a tail fairing interface 132. Thetail fairing interface 132 is configured to receive commands from theFCC 124 to control one or more actuators, such as a mechanical-hydraulicor electric actuators, for the active elevator 43 and/or active rudders45 of FIG. 2. In an embodiment, inputs 126 include an elevator pitchrate command for the tail fairing interface 132 to drive the one or moreactuators for pitch control of the active elevators 43 of FIG. 2. In anembodiment, inputs 126 include a rudder command for the tail fairinginterface 132 to drive the one or more actuators for positional controlof the active rudders 45 of FIG. 2.

Flight control system 120 may include an engine interface 133. Theengine interface 133 is configured to receive commands from the FCC 124to control engine(s) 24. In an embodiment, inputs 126 include a throttlecommand from the pilot to adjust the RPM of engine(s) 24. FCC 124 mayalso send commands to engine interface 133 to control the engine(s) incertain predefined operating modes (e.g., quiet mode).

The FCC 124 includes a processing system 134 that applies models andcontrol laws to augment commands based on aircraft state data. Theprocessing system 134 includes processing circuitry 136, memory 138, andan input/output (I/O) interface 140. The processing circuitry 136 may beany type or combination of computer processors, such as amicroprocessor, microcontroller, digital signal processor, applicationspecific integrated circuit, programmable logic device, and/or fieldprogrammable gate array, and is generally referred to as centralprocessing unit (CPU) 136. The memory 138 can include volatile andnon-volatile memory, such as random access memory (RAM), read onlymemory (ROM), or other electronic, optical, magnetic, or any othercomputer readable storage medium onto which data and control logic asdescribed herein are stored. Therefore, the memory 138 is a tangiblestorage medium where instructions executable by the processing circuitry136 are embodied in a non-transitory form. The I/O interface 140 caninclude a variety of input interfaces, output interfaces, communicationinterfaces and support circuitry to acquire data from the sensors 122,inputs 126, and other sources (not depicted) and communicate with therotor interface 128, the translation thrust interface 130, tail faringinterface 132, engine interface 133, and other subsystems (notdepicted).

In exemplary embodiments, the rotor interface 128, under control of theFCC 124, can control the upper rotor assembly 28 and lower rotorassembly 32 to pitch in different magnitudes and/or different directionsat the same time. This includes differential collective, where the upperrotor assembly 28 has a collective pitch different than the collectivepitch of the lower rotor assembly 32, in magnitude and/or direction.Differential pitch control also includes differential cyclic pitchcontrol, where the upper rotor assembly 28 has a cyclic pitch differentthan the cyclic pitch of the lower rotor assembly 32, in magnitude, axisof orientation (e.g., longitudinal or lateral) and/or direction. Thedifferential collective and the differential cyclic pitch control may beaccomplished using independently controlled swashplates in the upperrotor assembly 28 and lower rotor assembly 32. Alternatively,differential collective and the differential cyclic pitch control may beaccomplished using individual blade control in the upper rotor assembly28 and lower rotor assembly 32.

The ability to independently control the pitch of the upper rotorassembly 28 and lower rotor assembly 32 allows the lower rotor assembly32 to be adjusted due to its position beneath the upper rotor assembly28. The lower rotor assembly 32 is located in the downwash of the upperrotor assembly 28. To accommodate for this, the lower rotor assembly 32may have a collective pitch that differs from the collective pitch ofthe upper rotor assembly 28.

In the case of traditional helicopters, as the forward velocity of theaircraft increases, the velocity of the retreating blade relative to theairflow decreases. This causes a stall region to arise at the root ofthe retreating blade and expand towards to distal end of the blade asspeed increases. As this stall region increases, the overall lift vectorof the aircraft shifts from the center of the aircraft towards theadvancing blade which is providing the majority of lift for theaircraft. This imbalance of lift creates an unstable rolling moment onthe aircraft which is stabilized by a combination of reducing forwardflight and blade flapping, which reduces overall aircraft lift. With adual rotor aircraft, such as aircraft 10, the counter rotating rotorheads balance out the torque generated by each rotor head and alsobalances the lift generated by each advancing blade without the need forblade flapping or reducing the speed of the aircraft, or the need for awing. This is made possible by the rigid rotor system. Rigid rotorsallow for a reduced spacing between rotors. With two rigid rotors, theroll moments cancel at the main rotor shaft. Other rotor systems cangenerate opposing head moments, however, a greater spacing is requiredbetween rotors of those systems.

The use of upper rotor assembly 28 and lower rotor assembly 32 allowsthe pre-cone angle to be set on each individual rotor to reduce bendingstress on the blades. In a hinged rotor design, the hinges willnaturally go to an angle to reduce bending stress. On a rigid rotoraircraft, such as aircraft 10, there is no hinge, so the pre-cone angleis set to avoid the extra stress attributed to the bending moment. Auseful pre-cone angle is one where the centrifugal force of the bladepulling out matches the lift of the blade up. Due to the independentnature of the upper rotor assembly 28 and lower rotor assembly 32,differential pre-cone is used in aircraft 10. Differential pre-conerefers to the fact that the upper rotor assembly 28 and lower rotorassembly 32 have different pre-cone angles. The different pre-coneangles for the upper rotor assembly 28 and lower rotor assembly 32 helpmaintain tip clearance. In an exemplary embodiment, the pre-angle on theupper rotor assembly 28 is about 3 degrees and the pre-cone angle on thelower rotor assembly 32 is about 2 degrees.

Aircraft 10 is operational in a variety of modes, including take-off,cruise, landing, etc. Cruise mode refers to generally horizontal flight.During cruise, aircraft 10 can reach speeds of above about 200 knots,with speed reaching up to about 250 knots. During cruise mode, the mainrotor assembly 18 provides the majority of lift for the aircraft. Inexemplary embodiments and flight modes, the main rotor assembly 18provides greater than about 85% of the lift during cruise mode.

Aircraft 10 may assume various acoustic modes, depending on the flightstate. FCC 124 may control RPM of engines 24, RPM of propeller 42, andgearbox 90 to engage or disengage the propeller 42 to assume differentnoise levels. For example, at take-off noise may not be a concern, andthere would be no changes in aircraft operation to adjust the noiselevel. However, during engine start-up, the gearbox 90 may be disengagedsuch that the propeller 42 is decoupled from the main rotor system 18 toimprove ground safety. As the aircraft approaches a target, it may bedesirable to disengage the propeller 42 from the main rotor assembly 18using gearbox 90 and/or reduce RPM of engines 24 to reduce the noiseproduced by aircraft 10. The propeller 42 may be disengaged at variousother flight states (e.g., low speed cruise) to reduce noise. The RPM ofthe main rotor assembly 18 and RPM of propeller 42 may be independentlycontrolled (e.g., through gearbox 90 or FCC 124). This allows a varietyof flight states to be achieved.

The pilot may enter separate commands to reduce aircraft noise, forexample, disengaging the propeller 42, reducing engine RPM, andincreasing collective pitch as separate inputs. Alternatively, the pilotmay select a reduced noise mode (e.g., quiet mode) through single input,and the FCC 124 controls the various aircraft interfaces to achieve thedesired mode. For example, the pilot may select a reduced noise mode atinput 126, and the FCC automatically disengages the propeller 42,reduces the engine 24 RPM and/or increases collective pitch withoutfurther demand on the pilot.

The use of the translational thrust system 40 allows the aircraft 10 tomove forward or rearward (depending on the pitch of the propellerblades) independent of the pitch attitude of the aircraft. Cyclic isused to adjust the pitch attitude (nose up, nose down or level) of theaircraft while the translational thrust system 40 provides forward andrearward thrust.

The motor rotor assembly 18 system and the translational thrust system40 are connected through the main gearbox 26. A gear ratio of maingearbox 26 is selected so as to keep propeller 42 at a high efficiencyand suitable noise level during cruise mode. The gear ratio of maingearbox 26 dictates the ratio of the rotor speed of main rotor assembly18 to propeller speed of propeller 42.

Embodiments of aircraft 10 provide the pilot with increased situationalawareness by allowing the aircraft attitude (e.g., the angle oflongitudinal axis, L, relative to horizontal) to be adjusted by cyclicpitch of the main rotor assembly 18 and the forward and rearward thrustto be controlled by the translational thrust system 40. This allows avariety of flight modes to be achieved, which allows the pilot to bemore aware of their surroundings. Aircraft 10 can take off at ahorizontal attitude (e.g., axis L is horizontal), which also may bereferred to as vertical take-off Aircraft 10 may also fly forward orcruise with the nose angled upwards, nose angled downwards or level.Aircraft 10 can hover with the nose angled upwards or downwards orlevel. FIGS. 10 and 11 illustrate force vectors from the main rotorassembly and propeller for hover nose up and hover nose down,respectively. Aircraft 10 can also land substantially parallel to anon-horizontal or sloped surface by adjusting the attitude of theaircraft using cyclic of the main rotor assembly 18. The use of mainrotor assembly 18 for aircraft attitude and the translational thrustsystem 40 for thrust allows aircraft 10 to assume a variety of trimstates.

Embodiments provide independent control of the active elevators 43and/or active rudders 45 as controllable surfaces in the tail section41. The elevator surfaces 43 may be controlled independently by the FCC124 through the tail faring interface 132. The rudder surfaces 45 may becontrolled independently by the FCC 124 through the tail faringinterface 132.

The configuration of aircraft 10 and the controlled afforded by FCC 124allows aircraft 10 to provide a high bank angle capability at highspeeds. For example, in an exemplary embodiment, aircraft 10 can achievea bank angle of about 60 degrees at about 210 knots.

Aircraft 10 may make use of longitudinal lift offset in trim tocompensate for rotor-on-rotor aerodynamic interaction between the upperrotor assembly 28 and lower rotor assembly 32. Aircraft 10 may adjustdifferential longitudinal cyclic as a function of operational states ofthe aircraft (e.g., take-off, cruise, land, etc.). Differentiallongitudinal cyclic refers to upper rotor assembly 28 and lower rotorassembly 32 having different cyclic pitch along the longitudinal axis ofthe aircraft. Differential longitudinal cyclic may also be used togenerate yaw moments. Lift offset may be used to control aircraft, wherelateral lift offset adjusts roll and longitudinal lift offset adjustspitch.

FCC 124 may control RPM of engine(s) 24, RPM of propeller 42, andgearbox 90 to engage or disengage the propeller 42 to assume differentnoise levels. For example, at take-off noise may not be a concern, andthere would be no changes in aircraft operation to adjust the noiselevel. As the aircraft approaches a target, it may be desirable todisengage the propeller 42 using gearbox 90 and/or reduce RPM of engines24 to reduce the noise produced by aircraft 10. The propeller 42 may bedisengaged at various other flight states (e.g., high speed) to reducenoise. The RPM of the main rotor assembly 18 and RPM of propeller 42 maybe independently controlled (e.g., through gearbox 90).

The pilot may enter separate commands to reduce aircraft noise, forexample, disengaging the propeller 42 and reducing engine RPM asseparate inputs. Alternatively, the pilot may select a reduced noisemode (e.g., quiet mode) through single input, and the FCC 124 controlsthe various aircraft interfaces to achieve the desired mode. Forexample, the pilot may select a reduced noise mode at input 126, and theFCC automatically disengages the propeller 42 and/or reduces the engine24 RPM without further demand on the pilot.

Aircraft 10 provides the ability to approach a target and reverse thrustwhile maintaining an attitude directed at the target. FIG. 7 depictsaircraft 10 approaching a target 200. In a first state, 202, theaircraft 10 alters the pitch of blades 47 in propeller 42 to providereverse thrust to bring the aircraft to a quick stop. At state 204, themain rotor assembly 18 and propeller 42 are controlled to pitch aircraft10 towards target 200. At state 206, the propeller 42 is used to providereverse thrust to move away from target 200 and climb, while stillmaintaining an attitude with the nose of aircraft 10 facing target 200.

The use of a dual rotor system and translational thrust allows aircraft10 to eliminate the need for a variable angle between the main axis ofrotation of the rotor system (e.g., axis A in FIG. 1) and aircraftlongitudinal axis L. In conventional helicopters, the angle between themain axis of rotation of the rotor system and the aircraft longitudinalaxis L varies. This is due to the fact that conventional helicopterslack a translational thrust system 40 for use during cruise mode, orforward flight. In a conventional helicopter, forward flight is providedthrough cyclic pitch, which causes the aircraft to point nose down. Asthis nose down orientation is undesirable beyond a certain angle, theangle between the main axis of rotation of the rotor system and theaircraft longitudinal axis L is adjusted to bring the nose upwards,while still in forward flight.

By contrast, aircraft 10, with translational thrust system 40, does notneed to adjust the angle between the main axis of rotation of the rotorsystem (e.g., axis A in FIG. 1) and aircraft longitudinal axis L. Theangle between the main axis of rotation of the rotor system (e.g., axisA in FIG. 1) and aircraft longitudinal axis L for aircraft 10 remainsfixed during all flight modes, including take-off, cruise, landing, etc,unless otherwise commanded by a pilot of the aircraft 10.

As shown in FIG. 1, the rotor assembly 18 includes a rotor hub fairing37 generally located between and around the upper and lower rotorassemblies such that the rotor hubs 38 are at least partially containedtherein. The rotor hub fairing 37 provides drag reduction. Referring toFIG. 5, an exemplary rotor hub fairing 37 is shown. Rotor hub fairing 37is illustrated having generally elliptical, in cross-section, upper andlower hub fairings 111 and 112, and an airfoil-type shape (in horizontalcross-section) for the shaft fairing 103. The airfoil shape of the shaftfairing 103 includes a leading edge 114, and a trailing edge 115 aft ofthe upper and lower fairings 111, 112. The airfoil shape of the shaftfairing 103 additionally includes a chord (not shown) that connects theleading and trailing edges 114, 115 of the airfoil. In one embodiment,the airfoil shape, including the upper surface 116 and the lower surface117, is symmetrical about a plane extending along the length of theshaft fairing 103 and containing the axis of rotation, A. As notedabove, the upper and lower rotor hubs 38 may be positioned, at leastpartially, in the upper and lower fairings 111, 112.

The rotor hub fairing 37 is a sealed fairing, meaning there are few orno passages for air to travel through the interior of the rotor hubfairing 37. In conventional designs, control devices such as pushrods,are exposed near the rotor hubs. The surfaces of these componentsincrease drag on the rotor assembly. The air gaps between various rotorstructures (e.g., pushrods and main rotor shaft) also form areas ofdrag. The sealed rotor hub fairing 37 eliminates air pathways throughthe rotor hub structure, and eliminates drag associated with such airpaths.

Another feature to reduce drag on the rotor hub is positioning controlrods, such as push rods for rotor control, internal to the main rotorshaft. Referring to FIG. 3, pushrods for swashplates in the upper rotorassembly 28 and lower rotor assembly 32 are located internal to thelower rotor shaft 46 and upper rotor shaft 50. This prevents thepushrods from being exposed and increasing drag on the rotor hub. Theuse of a rigid rotor system aids in sealing the rotor hub faring 37.

In an exemplary embodiment, the distance between the hub of the upperrotor assembly 28 and the hub of the lower rotor assembly 32 ranges fromabout 2 feet to about 2.5 feet. In another exemplary embodiment, thedistance between the hub of the upper rotor assembly 28 and the hub ofthe lower rotor assembly 32 ranges from about 2.1 feet to about 2.4feet. In another exemplary embodiment, the distance between the hub ofthe upper rotor assembly 28 and the hub of the lower rotor assembly 32is about 2.29 feet (0.7 meters).

Aircraft 10 may employ an active vibration control (AVC) system toreduce vibration in the airframe 12. The use of a dual rotor, rigidrotor system tends to produce significant vibration in the airframe 12and its systems. FIG. 9 depicts an AVC system in an exemplaryembodiment. An AVC controller 300 executes an AVC control process toreduce vibration in aircraft 10. AVC controller 300 may be implementedas part of flight control system 120, executed by FCC 124, or may be aseparate controller. One or more sensors 302 are located in aircraft 10to detect vibration. Sensors may be located in a wide variety ofpositions, including airframe 12, gearbox 26, tail section 14, on mainrotor assembly 18, cockpit, etc. It is understood that these locationsare exemplary, and the AVC sensors 302 may be located in any position.AVC actuators 304 generate a force to dampen vibration in aircraft 10,as known in the art. AVC actuators 304 may be located in any position inthe aircraft.

In operation, AVC controller 300 receives vibration signals from the AVCsensors 302. AVC controller 300 provides control signals to the AVCactuators 304 to generate forces to reduce the vibration sensed by theAVC sensors 302. Control signals to the AVC actuators 304 may vary inmagnitude and frequency to cancel vibrations in aircraft 10. In anexemplary embodiment, AVC controller 300 operates in a feedback mode,where the control signals to AVC actuators 304 are adjusted in responseto measured vibration from AVC sensors 302. In an alternate embodiment,AVC controller 300 does not actively measure vibration through AVCsensors 302. Rather, the AVC controller 300 obtains the rotor speed(e.g., through an RPM signal) and applies a control signal to the AVCactuators 304, in an open loop control mode.

The use of independently controlled upper rotor assembly 28 and thelower rotor assembly 32, along with other control surfaces, provides theability to control yaw using a variety of elements. For example, below afirst speed, (e.g., 40 knots), the FCC 124 uses differential collectivepitch for yaw control. Above the first speed but below a second speed(e.g., 80 knots), a mix of differential collective and differentialcyclic may be used to control yaw. The differential cyclic may beapplied along the longitudinal and/or lateral axes of the aircraft.Further, wind direction may be measured by a sensor 122 and used toadjust the differential cyclic about the longitudinal and/or lateralaxes. Above the second speed (e.g., 80 knots), the active rudders 45 areused as controllable surfaces to control yaw. The FCC 124 providescommands to the tail fairing interface 132 to control the rudders 45 toadjust yaw.

The use of active elevator 43, with independent control of a leftelevator section and a right elevator section, provides for improvedstability control. Flight control system 120 performs mixing ofcollective pitch of main rotor assembly 18 and an angle of elevator 43to provide stability augmentation.

Embodiments may use wireless techniques to provide tip clearancemeasurements. FIG. 6A depicts a blade proximity monitoring system in anexemplary embodiment. At least one upper rotor blade and at least onelower rotor blade is equipped with at least one antenna 502. Antennas502 may be electric field antennas or magnetic field antennas. Antennas502 may be implemented using compact ferrite core or small diametermagnet wire in the form of coils around the blade spar or embedded inthe plane of the blade skin. The antennas 502 interact through the nearfield effect.

An oscillator 504 sends an excitation signal (e.g., 40 KHz) to a firstantenna 502L. The oscillator 504 sends another signal, which is to beused as feedback to monitor excitation signal status, to the bladeproximity monitor 508. It is understood that the excitation signal maybe sent to a plurality of antennas in different blades, includingmultiple antennas in the same blade. As the blades cross, a secondantenna 502U receives a signal emitted by the first antenna 502L. Areceiver 518 measures the magnitude of the excitation signal emittedfrom the first antenna 502L.

A blade proximity monitor 508 (e.g., a processor implemented controller)is mounted in the rotating system, e.g., in a rotor hub. This eliminatesnoise that may be introduced through a conventional slip ring used toconvey signals from a rotating system to a stationary system. The bladeproximity monitor 508 receives the conditioned output signal from thereceiver 518. Output signal from the second antenna 502U may befiltered, amplified or otherwise conditioned by the received 518. Theblade proximity monitor 508 also receives an RPM signal of the mainrotor assembly 18 from a contactor 510. Based on the magnitude of theexcitation signal applied to the first antenna 502L and the magnitude ofthe output signal from the second antenna 502U, blade proximity monitor508 can detect the distance between the first antenna 502L and thesecond antenna 502U. This provides an indication of the distance betweenthe rotor blades. The larger the magnitude of the output signal fromsecond antenna 502U, the closer the blades 36. The blade proximitymonitor 508 is also capable of determining the azimuth location of theblades 36 where the distance between the blades 36 is being measured.

The blade proximity monitor 508 may output the measured distance betweenthe blades to a rotor track and balance unit 512. The blade proximitymonitor 508 may output the measured distance between the blades and atwhat point in the azimuth that distance occurs to instrument system 514and to a pilot display 516. If the measured distance goes below athreshold, then an alert may be generated to the pilot that the bladesof the upper rotor assembly 32 and the lower rotor assembly 28 are tooclose to each other.

The use of a dual rotor, main rotor assembly 18 allows improvements incontrol of main rotor assembly 18. Flight control system 120 may applydifferent control envelopes to the upper rotor assembly 28 and the lowerrotor assembly 32. Flight control system 120 may impose differentcontrol ranges on the upper rotor assembly 28 and the lower rotorassembly 32 including control elements such as prioritization, gain vs.differential, collective versus cyclic, etc. The upper rotor assembly 28and the lower rotor assembly 32 may be independently controlled throughthe use of separate upper and lower swashplates. Alternatively, theupper rotor assembly 28 and the lower rotor assembly 32 may beindependently controller using individual blade control (IBC)techniques.

Aircraft 10 employs a fly-by-wire (FBW) control system to reduce pilotwork load. In an exemplary embodiment, FCC 124 determines the aircraftairspeed based on one or more sensors 122. The FCC 124 then adjusts thecollective pitch of the upper rotor assembly 28 and/or the lower rotorassembly 32 in response to the airspeed. FCC 124 may use a look up tablethat indexes airspeed to collective pitch. Alternatively, FCC 124 mayuse an algorithm to compute the collective pitch based on airspeed. Asnoted above, the collective pitch of upper rotor assembly 28 and thelower rotor assembly 32 may be the same or different.

Another feature to reduce pilot workload includes automaticallyadjusting the RPM and/or pitch of propeller 42 in response to a velocityor acceleration command from the pilot. Conventional systems wouldrequire the pilot to adjust propeller RPM and/or pitch throughindividual inputs. The flight control system 120 allows the pilot toenter a desired velocity or an acceleration, and the FCC 124 generatesthe proper commands to the translational thrust interface 130 toestablish an RPM and/or pitch to meet the desired velocity oracceleration.

In exemplary embodiments, the flight control system 120 controls themain rotor assembly 18 to prevent the tips of rotor blades 36 fromexceeding a threshold speed. In exemplary embodiments, the thresholdspeed may be Mach 0.9. This threshold would prevent the rotor blade tipsfrom exceeding the speed of sound. The threshold speed may vary, and maybe set to limit drag on the rotor blades to below a certain level. Inone embodiment, the FCC 124 determines air temperature from sensors 122.FCC 124 may also determine prevailing wind speed and direction fromsensors 122. The FCC 124 then computes the threshold speed based on thespeed of sound (e.g., Mach 1) at the sensed air temperature. The FCC 124may set the threshold to Mach 0.9, for example. FCC 124 then controlsRPM of the main rotor assembly 18 to prevent the rotor blade tips fromexceeding the threshold. In an exemplary embodiment, the FCC maintain85% of the nominal rotor RPM. FCC 124 may take into account prevailingwind direction and speed in controlling the RPM of the main rotorassembly 18. The Mach 0.9 threshold is only one example, and other speedthresholds may be employed to achieve desired results (e.g., reducedrag).

In exemplary embodiments, active elevator 43 is configured andcontrolled to improve stability be compensating for forces such aspropeller torque and/or rotor downwash. Elevator 43 includes a leftelevator and a right elevator on opposite sides of the axis of rotationof the propeller 42. The left elevator and right elevator may beindependently controlled to assume different positions. The tail fairinginterface 132 is configured to receive commands from the FCC 124 tocontrol one or more actuators, such as a mechanical-hydraulic orelectric actuators, to position the left elevator and right elevatorindependently. This independent control of the left elevator and rightelevator aids in compensating propeller torque and/or rotor downwash.

The left elevator and right elevator may also have different physicalconfigurations to compensate for compensating propeller torque and/orrotor downwash. The left elevator and right elevator may be offsetrelative to each other along the longitudinal and/or lateral axes ofaircraft 10. Further, the left elevator and right elevator may havedifferent geometries where one of the left elevator and right elevatoris larger than the other along the longitudinal and/or lateral axes ofaircraft 10. The left elevator and right elevator may have differingaerodynamic surfaces (e.g., airfoils) as well.

The cockpit of aircraft includes a single, semi-active, collective input(e.g., stick) positioned between the two pilot seats.

Exemplary embodiments of aircraft 10 provide a much smaller footprintthan existing aircraft. This makes aircraft 10 well suited for missionsin confined terrain, urban settings, and shipboard operations. FIG. 8presents front, side and top views of an exemplary aircraft. One featurecontributing to the reduced footprint is the location of the main rotorshaft relative to the airframe 12. As shown in FIG. 1, the axis ofrotation A, of the main rotor assembly 18, intersects longitudinal axis,L, along a span of axis L, extending from the nose of the aircraft tothe tip of the hub of propeller 42. In an exemplary embodiment, the axisof rotation A is located at about a 44% station (STA) of the fuselage orairframe 12.

In an exemplary embodiment, there is about 5.2 inches from the mainrotor pylon to the blade hub centerline. In an exemplary embodiment,there is about 0.7 inch hub clearance to the main rotor pylon. In anexemplary embodiment, the rotor blades 36 extend beyond the nose of theaircraft by about 13 inches (0.33 meters). In an exemplary embodiment,rotor blades 36 extend beyond the nose of the aircraft by about 6.9% ofthe blade span, which may be about 188 inches.

The use of a rigid rotor system, along with the rotor shaft position(e.g., axis A) allows for much easier air-to-air refueling. The stiffrotor blades 36 ease air-to-air refueling by reducing blade flapping,which may result in a blade contacting a tanker fuel line duringrefueling.

Aircraft 10 provides an improved glide slope angle of about 5-to-1 toabout 6-to-1. This is due to the propeller 42 taking energy out of theairstream, inputting energy into the gear box 26 to increase the speedof the main rotor assembly 18 during autorotation. As shown in FIGS. 3and 4, the main gear box 26 interconnects the main rotor assembly 18 andpropeller 42. During autorotation, the airflow rotates propeller 42,which will subsequently rotate the main rotor assembly 18 and thusincrease lift. Propeller 42 also helps stabilize aircraft 10 duringdecent by acting like a parachute and a rudder, both slowing aircraft 10and helping to direct aircraft 10 to maintain control. The ability tofly aircraft 10 in a nose down attitude also improves glide slope angle.

In an exemplary embodiment, the distance between the hub of the upperrotor assembly 28 and the hub of the lower rotor assembly 32 ranges fromabout 2 feet to about 2.5 feet. In another exemplary embodiment, thedistance between the hub of the upper rotor assembly 28 and the hub ofthe lower rotor assembly 32 ranges from about 2.1 feet to about 2.4feet. In another exemplary embodiment, the distance between the hub ofthe upper rotor assembly 28 and the hub of the lower rotor assembly 32is about 2.29 feet. In another exemplary embodiment, the distancebetween a midpoint of a blade in the upper rotor assembly 28 and amidpoint of a blade in the lower rotor assembly 32 is about 29.0 inches.In another exemplary embodiment, the distance between a tip of a bladein the upper rotor assembly 28 and a tip of a blade in the lower rotorassembly 32 is about 31.0 inches. In another exemplary embodiment, thedistance between the hub of the upper rotor assembly 28 and the hub ofthe lower rotor assembly 32 is about 14% of the blade span, which may beabout 188 inches.

The terminology used herein is for the purpose of describing particularembodiments only and is not intended to be limiting of the invention.While the description of the present invention has been presented forpurposes of illustration and description, it is not intended to beexhaustive or limited to the invention in the form disclosed. Manymodifications, variations, alterations, substitutions, or equivalentarrangement not hereto described will be apparent to those of ordinaryskill in the art without departing from the scope and spirit of theinvention. Additionally, while the various embodiment of the inventionhave been described, it is to be understood that aspects of theinvention may include only some of the described embodiments.Accordingly, the invention is not to be seen as limited by the foregoingdescription.

1. An aircraft comprising: an airframe; an extending tail; a counterrotating, coaxial main rotor assembly including an upper rotor assemblywith an upper blade and a lower rotor assembly with a lower blade; afirst antenna in one of upper blade and the lower blade, and a secondantenna in the other of the upper blade and the lower blade; anoscillator to apply an excitation signal to the first antenna; and ablade proximity monitor to monitor a magnitude of the excitation signaland an output signal from the second antenna to determine a distancebetween the upper blade and the lower blade.
 2. The aircraft of claim 1wherein: the blade proximity monitor is mounted in a rotating portion ofthe main rotor assembly.
 3. The aircraft of claim 1 wherein: the bladeproximity monitor is configured to determine the azimuth location of theblades where the distance between the upper blade and the lower blade isdetermined.
 4. The aircraft of claim 3 wherein: the blade proximitymonitor is configured to transmit at least one of the distance betweenthe upper blade and the lower blade and the azimuth location to at leastone of an instrument system and a pilot display.
 5. The aircraft ofclaim 1 wherein: the blade proximity monitor is configured to receive anRPM signal of the main rotor assembly from a contactor.
 6. The aircraftof claim 1 wherein: the antennas are electric field antennas.
 7. Theaircraft of claim 1 wherein: the antennas are magnetic field antennas.8. The aircraft of claim 1 wherein: the antennas are composed of aferrite core material in the form of coils around the blade spar orembedded in the plane of the blade skin.
 9. The aircraft of claim 1wherein: the antennas are composed of a small diameter magnet wire inthe form of coils around the blade spar or embedded in the plane of theblade skin.
 10. The aircraft of claim 1 wherein: the blades are rigidlyconnected to the upper rotor assembly and the lower rotor assembly. 11.A method of operating an aircraft, comprising: generating, using anoscillator, an excitation signal; transmitting, the excitation signalthrough a first antenna in one of an upper blade and a lower blade;receiving, the excitation signal through a second antenna in the otherof the upper blade and the lower blade and transmitting an outputsignal; and determining, with a blade proximity monitor, the distancebetween the upper blade and lower blade by using the magnitude of theexcitation signal and an output signal from the second antenna.
 12. Themethod according to claim 11 wherein: the blade proximity monitor ismounted in a rotating portion of the main rotor assembly.
 13. The methodaccording to claim 11 wherein: the blade proximity monitor determinesthe azimuth location of the blades where the distance between the upperblade and the lower blade is determined.
 14. The method according toclaim 13 wherein: the blade proximity monitor transmits at least one ofthe distance between the upper blade and the lower blade and the azimuthlocation to at least one of an instrument system and a pilot display.15. The method according to claim 11 wherein: the blade proximitymonitor receives an RPM signal of the main rotor assembly from acontactor.
 16. The method according to claim 11 wherein: the antennasare electric field antennas.
 17. The method according to claim 11wherein: the antennas are magnetic field antennas.
 18. The methodaccording to claim 11 wherein: the antennas are composed of a ferritecore material in the form of coils around the blade spar or embedded inthe plane of the blade skin.
 19. The method according to claim 11wherein: the antennas are composed of a small diameter magnet wire inthe form of coils around the blade spar or embedded in the plane of theblade skin.
 20. The method of claim 11 wherein: the blades are rigidlyconnected to the upper rotor assembly and the lower rotor assembly.